Method of forming composite wound structure

ABSTRACT

An elongate hollow structure with a plurality of reinforcements arranged in a pattern and acting as spacers between a pair of skins. The skins are of resin impregnated wound filaments with reinforcing strips of facing preimpregnated wound filaments that crisscross at the spacing reinforcements, abutting panels extend between adjacent reinforcing spacers to fill the space between the skins, and the structure is bonded with resin into a composite structure.

This is a division of application Ser. No. 930,457 filed Aug. 2, 1978now U.S. Pat. No. 4,230,293.

BACKGROUND OF THE INVENTION

Present day airplanes utilize an efficient structure that is lightweightfor a given loading. Composite materials are recognized as offering apotential for an even more efficient, weight wise, structure. In U.S.Pat. No. 2,817,484 to Stenzel it shows a fuselage type structure withspirally wound hollow metal members, and longitudinal members bothbonded together with a bonding agent.

SUMMARY OF THE INVENTION

An inner and an outer skin of filaments wound circumferentially andlongitudinally each have facing crisscrossing helically wound filamentsin reinforcing strips. The skins are spaced apart with reinforcing plugslocated to contact the strips at the intersections. The plugs haveinward extending fastener sockets and the inner skin and reinforcingstrips have contiguous openings to permit entry into the sockets. Panelsextend between adjacent plugs and the two skins and all the componentsare integrally joined with resins into a composite structure. Frames forsupport of openings through the structure are integrally located withinthe composite.

DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a fragmented perspective view of the composite structure ofthis invention with parts broken away to show the buildup of thecomposite.

FIG. 2 is a blown up partial plan view of a portion of the structure ofFIG. 1.

FIG. 3 is a plan view of a panel taken from FIG. 2.

FIG. 4 is a sectional view taken along lines 4--4 of FIG. 2.

FIG. 5 is a sectional view taken along lines 5--5 of FIG. 3.

FIG. 6 is a sectional view taken along lines 6--6 of FIG. 2.

FIG. 7 is a sectionalized plan view of a different embodiment of thepanel of FIG. 3.

FIG. 8 is a view taken along lines 8--8 of FIG. 7.

FIG. 9 is a perspective view of a spacer used in this invention.

FIG. 10 is a plan view showing framing for an opening through thestructure of this invention.

DETAILED DESCRIPTION

An elongate hollow composite structure 10 made up of an inner skin 12and an outer skin 14 each made up of filaments which preferably arewound at ±10° or essentially longitudinally and others wound at ±80° oressentially circumferentially. These fibers may be any of the highstrength fibers such as glass, boron, graphite or kelvar with graphitepreferred, and may be in the form of individual filaments, a group offilaments or in a tape. The filament are embedded in a plastic or resinsuch as an epoxy, a polyamide or a polyimide which may be preimpregnatedonto the filaments or may be applied to the wound fibers. A cap orreinforcing strip 16 in a spaced apart pattern such as a geodesicpattern crosscross each other and are in contact with the inner skin,and a similar crisscrossing reinforcing strip 18 is in contact with theouter skin. The center of crossing of these reinforcing cap strips areradially aligned with respect to each other. These strips are each madeup of spirally wound filaments embedded in a resin with the same typesof fibers and resins as are used in the skins. A series of reinforcingplugs 20 are located to extend between the reinforcing strips with theaxis of the plugs extending radially between the intersection of thecrisscrossing strips. These plugs, as best shown in FIG. 9, arepreferably cruciform in shape with center section 22 and with fourtapered blades 24. A socket 26 is axially located in the center sectionand faces radially inward with respect to the composite structure toaccept a fastener such as the bolt 28 shown. The blades are positionedto extend in the direction of the reinforcing strips and are slightlyindented top and bottom at 30 to accommodate the extra thickness 17 and19 of the respective inner and outer reinforcing strips where they crosseach other. These reinforcing plugs or spaces may be of any light weighthigh strength material with chopped graphite fibers embedded in an epoxyresin preferred. Extending to and filling the space between the innerand outer skins with adjoining reinforcing strips are a series of coreblocks or panels 32 that are positioned at the corners with thereinforcing plugs and abut each other along the sides. These panels areof a light weight material such as honeycomb or closed cell foamedresins and are contoured to fit the space. These panels are best shownin FIGS. 2, 3 and 5 are shaped at the corners 34 to match the contour ofthe sides 36 of the reinforcing plugs 20, at bottom 38 and top 40 tomatch the contour of the skins, and have reduced thickness near thebottom edge 42 and top edge 44 to allow for the reinforcing strips 16and 18. These panels preferably also have resin impregnated filaments intapes with the filaments crisscrossing on a bias, or in other words at±45° with the tape 46 wound around the outer edge 48 and with the tapeof a width to overlap part way onto the sides at the bottom and topedges 42 and 44 to extend under the reinforcing strips. When the resinsin these positioned components are cured it makes up a compositestructure having inner 12 and outer 14 skins spaced apart with coremembers 32 and reinforcing plugs 20 and reinforced with crisscrossing Ibeam like members made up of the reinforcing plugs at the intersectionand the inner 16 and outer 18 reinforcing strips joined by reinforcingmembers 46.

A modified panel 32a is used to permit viewing through the structure. InFIGS. 7 and 8 the panel 32a is in two sections with each section cut outat 50 with shaped edges 52 to fit into a U-shaped portion 54 of awindowframe 56. When the two sections are joined around the windowframethe corners 34a, edges 42a and 44a, and resin impregnated biased tape46a will be shaped to fit between skins 12 and 14 and reinforcing strips16 and 18, and having corners located by the reinforcing plugs 20. Afterthe composite structure is formed the skins will be cut out around theinside circumference 58 of the windowframe and the skins removed toexpose a window opening 60. The transparent material for the window andthe details of mounting the same are not shown.

FIG. 10 shows the mounting for a doorframe 62 to provide an openingthrough the composite structure. The doorframe has a U-shaped outerperiphery 64 into which contoured edges 42b and 44b of core panels 32bextend. These edges are shown covered with resin impregnated biased tape46b. A door 65 has a structural member 66 located adjacent the insideperiphery 67 of the doorframe 62. This structural member has a U-shapedinner periphery 68 into which contoured edges of core panels extend.Skins 12 and 14 and reinforcing strips 16 and 18 extend over thedoorframe and over the door. Reinforcing plugs 20 are located at theintersection of the reinforcing strips. After the composite is cured theskins and reinforcing strips are cut through around the inside periphery67 of the doorframe to provide an opening through the structure and toprovide a door for the opening.

To prepare the composite a mandrel 70 which has an outside contour whichis the shape of the inside wall of the finished structure is used. Themandrel has a series of index pins 72 extending outward in a radialdirection and located to be the midpoints for the intersection ofreinforcing strips 16. A parting agent is placed over the mandrel andskin 12 is formed using resin impregnated fibers that are laid down in aspiral fashion to completely cover the mandrel with a skin of thedesired thickness. For many applications this skin will be of graphitefibers impregnated with epoxy resin, and built up to about 0.022" thick.During the layup of the skins the windings, which spirally windessentially horizontally and others spirally wind essentiallycircumferentially, settle around the pins so that the pins protrudethrough the windings.

Next the cap strips or reinforcing strips 16 of resin impregnatedfilaments are spirally wound in a spaced apart pattern with the stripsintersecting at and settling around the index pins 72. These strips mayextend longitudinally and circumferentially, however, it is preferredthey be wound at about 45 degrees and in a geodesic pattern and about0.055" thick.

In the next step the reinforcing spacers or plugs 20 which preferablyare cruciform in shape are placed over each of the index pins so thatthe blades 20 of the spacers extend in the direction of the reinforcingstrips 16 and the axis of the plugs extend radially. The index pinsextend into the inwardly directed sockets 26 to hold and position thereinforcing plugs. At least some of these sockets are threaded to accepta threaded fastener, however, the index pins insert into the position,but do not thread into the sockets.

The formed core panels 32 are then positioned with surface 38 againstthe skin 12 and the resin impregnated biased tape 46 that surrounds theedge of the panels contacting the sides 36 of contiguous plugs at thecorners and abutting each other elsewhere around the periphery. In thosepositions where windows are desired a windowframe 56 with special panels32a are used in place of the regular panels 32. In the locations wheredoors are desired, the doorframe 62 with adjacent panels 32b, and thedoors 65 with adjacent panels 32c replace some of the regular panels 32.

As the next step resin impregnated filaments are wound in strips 18 tocrisscross at the axis of the reinforcing plugs. These strips areoriented the same as the first layer of strips 16. A layer of resinimpregnated filaments is then laid in a crisscrossing fashion withfibers spirally wound essentially longitudinally and circumferentiallyto form an outer skin 14. The mandrel with laid up components is thenheated to cure the resins and form a composite structure. The corematerial expands somewhat at the curing temperature to provide pressureon the skins and reinforcing strips. Once the resins are cured themandrel with index pins is collapsed, the materials are cut away fromboth the inside and outside layers at the windowframes to expose thewindows and from around the doorframes to permit entry into and out ofthe composite structure.

We claim:
 1. A method of preparing a composite fuselage, the stepscomprising: utilizing a collapsible mandrel having a plurality ofoutwardly extending index pins, covering the mandrel with a layer ofresin impregnated filaments wound with fiber orientation ±10° and at±80° while setting the windings around the index pins so that the pinsprotrude through, reinforcing the covering with cap strips of resinimpregnated filaments wound with fiber orientation at about ±45° and thefilaments intersect at and settle around the index pins, placing arecessed spacer plug over each index pin, positioning contoured panelswith corners of each panel located by adjacent plugs and edges of thepanels abutting to cover the surface, winding cap strips with resinimpregnated filaments and intersecting the strips over each plug,covering with a layer of resin impregnated filaments wound with fiberorientation at ±10° and at ±80°, curing the resins to form a compositestructure, and removing the mandrel with indexing pins.
 2. A method ofpreparing a composite fuselage as in claim 1, steps further comprising:utilizing a cruciform shape for the plugs and positioning said plugswith blades to extend in the direction of the strips.
 3. A method ofpreparing a comosite fuselage is in claim 2, steps further comprising:threading the recess in at least some of the spacer plugs for acceptinga fastener.
 4. A method of preparing a composite fuselage as in claim 1,steps further comprising: wrapping a resin impregnated tape with biasedcrisscrossing filaments around the edges of the panels prior to placingthe panels in position, and selecting biased tape of a width to overlaponto the sides for extending under the cap strips.
 5. A method ofpreparing a composite fuselage as in claim 4, steps further comprising:mounting windowframes in some of the panels before placement of thepanels, and cutting away the inner and outer covering member forexposing windows.
 6. A method of preparing a composite fuselage as inclaim 4, steps further comprising: mounting doorframes in some of thepanels before placement of the panels and cutting away the inner andouter covering member for exposing the doorframes.
 7. A method ofpreparing a composite fuselage as in claim 4, steps further comprising:mounting doorframes in some of the panels before placement of the panelsand cutting away the inner and outer covering members for exposingdoorframe openings and doors within the openings.